The statements in this section merely provide background information related to the present disclosure and may not constitute prior art.
An aircraft is moved by several turbojet engines each housed within a nacelle. The propulsion unit constituted by a turbojet engine and the nacelle that receives it is shown in FIG. 1 referred thereto.
The propulsion unit 1 comprises a nacelle 3 supporting a turbojet engine 5. The propulsion unit 1 is connected to the aircraft fuselage (not shown) for example by means of a pylon 7 intended to be suspended under a wing of the aircraft.
The nacelle 5 generally has a tubular structure comprising an upstream section 9 defining an air inlet upstream of the turbojet engine 5, a median section 11 intended to surround a fan of the turbojet engine, a downstream section 13 comprising an outer cowling 15 able to accommodate a thrust reverser device and intended to surround the combustion chamber of the turbojet engine, and is generally terminated by an ejection nozzle whose outlet is located downstream of the turbojet engine.
This nacelle accommodates the turbojet engine 5 which can be of the bypass type, adapted to generate, via the blades of the rotating fan, a hot air flow (also called primary flow), coming from the combustion chamber of the turbojet engine, and a cold air flow (secondary flow) that circulates outside the turbojet engine through a flow path 17 (half-flow path 17a shown in FIG. 2), also called annular channel, formed between a fairing of the turbojet engine and an inner wall 18 (inner half-wall 18a shown in FIG. 2) of the outer structure 21 (outer half-structure 21a shown in FIG. 2) of the nacelle. The two air flows are ejected from the turbojet engine from the rear of the nacelle.
Reference is made to FIG. 2 showing a right half-shell 13a of a nacelle that constitutes, with a second half-shell (not represented, obtained by symmetry relative to a median plane of the nacelle), the downstream structure 13 of the nacelle able to surround the combustion chamber of the turbojet engine (not represented in this figure). It should be noted that this downstream structure can integrate a thrust reversal device, on the understanding that the invention also applies to a smooth nacelle case, that is to say devoid of a thrust reversal device.
The references FRONT and REAR respectively designate the front (upstream) and rear (downstream) portions of the half-shell 13a, relative to the direction of the air flow intended to circulate inside this half-shell 13a. 
In this case, this half-shell 13a includes an inner half-structure 19a, defining a half-cavity C intended to receive the turbojet engine (not represented). An inner structure 19 is obtained by the assembly of two inner half-structures 19a and 19b (only the half-structure 19a is shown in FIG. 2, the half-structure 19b being positioned symmetrically with the half-structure 19a relative to the median plane of the nacelle).
This half-shell 13a also includes an outer structure 21a defining, with the inner half-structure 19a, a half-flow path 17a intended to be traversed by a cold air flow circulating between the front and the back of the half-shell 13a and defining, with the half-flow path obtained by symmetry relative to the median plane of the nacelle, the flow path 17 or annular channel.
The connection of the engine to the aircraft is carried out by means of a support structure comprising two upper longitudinal half-beams 23a, 23b (only the half-beam 23a is shown in FIG. 2, the half-beam 23b being positioned symmetrically with the half-beam 23a relative to the median plane of the nacelle), conventionally called 12 o'clock beams because of their position at the top of the nacelle and two lower longitudinal half-beams 25a, 25b (only the half-beam 25a is shown in FIG. 2, the half-beam 25b being positioned symmetrically with the half-beam 25a relative to the median plane of the nacelle), conventionally called 6 o'clock beams because of their position in the lower portion of the nacelle. The lower “6 o'clock” half-beams 25a, 25b are conventionally faired by means of fairing sheets 26 (represented in FIG. 3 showing the nacelle 3 viewed from the bottom) intended to come into contact with the outer air flow flowing around the nacelle.
The 12 o'clock and 6 o'clock half-beams are interconnected, on the one hand, via the inner structure 19 surrounding the turbojet engine and, on the other hand, via a substantially annular structure called front frame and generally formed of two front half-frames 27a, 27b (only the front half-frame 27a is shown in FIG. 2, the front half-frame 27b being positioned symmetrically with the front half-frame 27a relative to the median plane of the nacelle) each extending between said corresponding half-beams on both side of the median plane of the nacelle. This front frame is intended to be fixed to the periphery of a downstream edge of a casing of the fan engine and thus contribute to the recovery and transmission of forces between the different portions of the nacelle and of the turbojet engine. Furthermore, in the case of a nacelle equipped with a cascade thrust reverser device, the front frame is also used to support the cascades of the thrust reverser.
Conventionally, a cascade thrust reverser comprises two half-cowls (forming the outer cowling 15 shown in FIG. 1) each slidably mounted on the upper 23a, 23b and lower 25a, 25b half-beams. For this purpose, the upper and lower half-beams are generally equipped with primary and secondary guide rails allowing a sliding movement of the half-cowls of the cascade thrust reverser, on its associated half-beam between alternately a position of the thrust reverser in direct jet according to which the half-cowls ensure the aerodynamic continuity of the nacelle and a position of the thrust reverser in reverse jet according to which the half-cowls are displaced downstream of the nacelle.
The bypass ratio of a turbojet engine is defined by the ratio between the air mass of the cold air flow passing through the flow path of the propulsion unit and the mass of the hot air flow passing through the turbojet engine. In engines with a high bypass ratio (for example a ratio of 10), the diameter of the flow path 17 of the cold air flow is increased relative to an engine with a lower bypass ratio.
The increase in the diameter of the flow path 17 results in a radial distance, relative to the longitudinal axis of the propulsion unit, of the lower “6 o'clock” half-beams 25a, 25b. 
Referring to FIGS. 3 to 5 on which is schematized, for a better understanding, this radial distance of the lower “6 o'clock” half-beams 25a, 25b induced by the increase in the diameter of the flow path 17.
The radial distance of the lower half-beams 25a, 25b results in a radial distance of the fairing sheets 26 (shown in FIG. 3 illustrating the nacelle viewed from the bottom) fixed on the outer wall of the lower “6 o'clock” half-beams and coming into contact with an outer air flow Fext flowing around the nacelle.
Referring to FIG. 4 illustrating the downstream section 13 of the nacelle in longitudinal section on which there is shown an aerodynamic surface 29 defined by the fairing sheets 26 and an aerodynamic surface 31 that would be obtained when the diameter of the flow path 17 would have been increased in order to obtain an engine with a higher bypass ratio.
It is noticed in this figure that the outer aerodynamic surface 31 of the nacelle has become radially more distant relative to the longitudinal axis 33 of the nacelle and relative to the outer aerodynamic surface 29 obtained for an engine with a lower bypass ratio.
This increase in the diameter of the nacelle results in an increase in the size and the mass of the nacelle. In addition, this also directly results in an increase in the size of the “beavertail” or “six o'clock rear beam fairing,” a term used to designate the fairing 35 in the form of a “beaver tail” downstream of the nacelle and shown in FIGS. 3 and 5. The increase in the size and the mass of the nacelle and in the “beavertail” causes an increase in the aerodynamic drag of the nacelle.